The present invention relates generally to gas turbine engines, and, more specifically, to heated components thereof.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through an annular turbine nozzle and cooperating turbine stages which extract energy therefrom. The high pressure turbine nozzle includes a row of hollow vanes mounted in arcuate outer and inner bands and directs the combustion gases into a first stage high pressure turbine.
The turbine includes a row of hollow rotor blades mounted on a supporting rotor disk which is joined to the compressor for powering thereof during operation. A low pressure turbine follows the high pressure turbine and typically includes multiple stages of turbine rotor blades which extract additional energy from the combustion gases for typically powering a fan disposed upstream from the compressor. In this turbofan configuration of the gas turbine engine, the fan produces propulsion thrust for powering an aircraft in flight.
Since the combustor and turbine components are directly exposed to the hot combustion gases, they are typically cooled during operation by diverting a portion of the pressurized air from the compressor for use as a coolant channeled through these various hot components. Any air diverted from use in the combustion process decreases the overall efficiency of the engine and should be minimized. However, the hot engine components must be suitably cooled for ensuring a suitable useful life thereof.
Various cooling configurations are common for the combustor liners, turbine nozzle vanes, nozzle bands, turbine blades, turbine shrouds, as well as for exhaust liners for both commercial and military engines. These various cooling configurations include dedicated cooling channels, with various forms of cooling holes through which the cooling air is re-introduced into the combustion gas flowpath. Inclined film cooling holes are common in various ones of these engine components and provide internal convection cooling of the walls thereof, while also forming a protective cooling film of air along the exposed or outer surface of the wall which faces the hot combustion gases.
For additional protection from the hot combustion gases, the exposed outer walls of the hot engine components may be covered with a thermal barrier coating (TBC) which provides thermal insulation. The benefits of thermal barrier coatings are well known, and include enhanced performance and life of the gas turbine engine.
Typical thermal barrier coatings include ceramic material commonly requiring a corresponding metallic bond coating or layer firstly adhered to the metal surface of the engine components for in turn providing an improved bond with the thermal barrier coating. The TBC and cooperating bond coat layer ensure a strong bond thereof with the underlying metal substrate for ensuring the durability of the TBC for extended life in engine operation. In alternate developments, thermal barrier coatings may be applied to the metal substrate without a bond coat.
However, the durability of the TBC is affected by the operational temperature of its underlying support including the metal substrate and bond coating when used. As the bond coating experiences elevated temperature during operation, it is subject to temperature degradation thereof, and degradation of the bond coating weakens the TBC/bond coating interface and leads in turn to TBC spallation.
The ability to cool both the bond coating and the TBC thereatop is limited by the cooling configurations of the underlying metal substrate or wall. The component walls are typically relatively thin and cooled from their inner or back sides by channeling the cooling air thereover. The walls may also include the inclined film cooling holes which provide local internal convection cooling at each of the holes, in addition to providing the protective cooling air over the exposed or external surface of the component covered by the TBC.
Accordingly, it is desired to provide an improved thermal barrier coating system with enhanced cooling for enhancing the durability and life of the TBC during operation.
A gas turbine engine component includes a perforate metal wall having pores extending therethrough. The wall has a first surface covered by a thermal barrier coating. The pores have first ends which are covered by the thermal barrier coating, and the pores are ventilated from an opposite second surface of the wall for cooling the thermal barrier coating.